The present invention generally relates to ceramic-based articles and processes for their production. More particularly, this invention is directed to ceramic-based articles produces to include metallic regions that define detailed features, for example, dovetails, shanks, platform features and tip shrouds of gas turbine airfoil components.
Higher operating temperatures for gas turbines are continuously sought in order to increase their efficiency. Though significant advances in high temperature capabilities have been achieved through formulation of iron, nickel and cobalt-base superalloys, alternative materials have been investigated. Ceramic materials are a notable example because their high temperature capabilities can significantly reduce cooling air requirements. As used herein, ceramic-based materials encompass homogeneous ceramic materials as well as ceramic matrix composite (CMC) materials. CMC materials generally comprise a ceramic fiber reinforcement material embedded in a ceramic matrix material. The reinforcement material may be discontinuous short fibers dispersed in the matrix material or continuous fibers or fiber bundles oriented within the matrix material. The reinforcement material serves as the load-bearing constituent of the CMC in the event of a matrix crack. In turn, the ceramic matrix protects the reinforcement material, maintains the orientation of its fibers, and serves to dissipate loads to the reinforcement material. Silicon-based composites, such as silicon carbide (SiC) as the matrix and/or reinforcement material, are of particular interest to high-temperature applications, for example, high-temperature components of gas turbines including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry.
Continuous fiber reinforced ceramic composites (CFCC) are a type of CMC that offers light weight, high strength, and high stiffness for a variety of high temperature load-bearing applications, including shrouds, combustor liners, vanes (nozzles), blades (buckets), and other high-temperature components of gas turbines. A notable example of a CFCC has been developed by the General Electric Company under the name HiPerComp®, and contains continuous silicon carbide fibers in a matrix of silicon carbide and elemental silicon or a silicon alloy. SiC fibers have also been used as a reinforcement material for a variety of other ceramic matrix materials, including titanium carbide (TiC), silicon nitride (Si3N4), and alumina (Al2O3).
Examples of CMC materials and particularly SiC/Si—SiC (fiber/matrix) CFCC materials and processes are disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. One such process is known as “prepreg” melt-infiltration (MI), which in general terms entails the fabrication of CMCs using multiple prepreg layers, each in the form of a tape-like structure comprising the desired reinforcement material and a precursor of the CMC matrix material, as well as one or more binders and typically carbon or a carbon source. The prepreg must undergo processing (including firing) to convert the precursor to the desired ceramic. Prepregs for CFCC materials frequently comprise a two-dimensional fiber array comprising a single layer of unidirectionally-aligned tows impregnated with a matrix precursor to create a generally two-dimensional laminate.
For purposes of discussion, a low pressure turbine (LPT) blade 10 of a gas turbine engine is represented in FIG. 1. The blade 10 is an example of a component that can be produced from a ceramic-based material, including CMC materials. The blade 10 is generally represented as being of a known type and adapted for mounting to a disk or rotor (not shown) within the turbine section of an aircraft gas turbine engine. For this reason, the blade 10 is represented as including a dovetail 12 for anchoring the blade 10 to a turbine disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented in FIG. 1, the interlocking features comprise protrusions referred to as tangs 14 that engage recesses defined by the dovetail slot. The blade 10 is further shown as having a platform 16 that separates an airfoil 18 from a shank 20 on which the dovetail 12 is defined. The blade 10 may be further equipped with a blade tip shroud (not shown) which, in combination with tip shrouds of adjacent blades within the same stage, defines a band around the blades that is capable of reducing blade vibrations and improving airflow characteristics. By incorporating a seal tooth, blade tip shrouds are further capable of increasing the efficiency of the turbine by reducing combustion gas leakage between the blade tips and a shroud surrounding the blade tips.
Because they are directly subjected to hot combustion gases during operation of the engine, the airfoil 18, platform 16 and tip shroud have very demanding material requirements. The platform 16 and blade tip shroud (if present) are further critical regions of a turbine blade in that they create the inner and outer flowpath surfaces for the hot gas path within the turbine section. In addition, the platform 16 creates a seal to prevent mixing of the hot combustion gases with lower temperature gases to which the shank 20, its dovetail 12 and the turbine disk are exposed, and the blade tip shroud is subjected to creep due to high strain loads and wear interactions between its seal tooth (if present) and the shroud surrounding the blade tips. The dovetail 12 is also a critical region in that it is subjected to wear and high loads resulting from its engagement with a dovetail slot and the high centrifugal loading generated by the blade 10.
Current state-of-the-art approaches for fabricating ceramic-based turbine blades have involved integrating the platform 16, dovetail 12, airfoil 18 and tip shroud (if present) as one piece during the manufacturing process, much like conventional investment casting techniques currently used to make metallic blades. However, the platform 16, dovetail 12, tangs 14 and tip shroud represent detailed geometric features of the blade 10 that pose substantial challenges to designing, manufacturing and integrating CMC components into an affordable, producible design for turbine applications. For example, the process of integrating a platform 16 and tip shroud with the airfoil 18 using CMC materials creates complexities in the design and manufacturing process, and can result in a process that can be too expensive to be economically practical. Furthermore, the platform 16, dovetail 12 and its tangs 14 have interface/support functions that can require structural interface capabilities that can be difficult to achieve with CMC materials. In addition, the low strain-to-failure capabilities of typical CMC materials and the possibility of undesirable wear interactions between tip shroud seal teeth and conventional shrouding materials pose additional challenges to implementing CMC materials in shrouded blade designs.